1. Field
The present invention relates to turbine airfoils for gas turbine engines, and in particular to a turbine airfoil having one or more inserts for near-wall cooling.
2. Description of the Related Art
In a turbomachine, such as an axial flow gas turbine engine, air is pressurized in a compressor section and then mixed with fuel and burned in a combustor section to generate hot combustion gases. The hot combustion gases are expanded within a turbine section of the engine where energy is extracted to power the compressor section and to produce useful work, such as turning a generator to produce electricity. The hot combustion gases travel through a series of turbine stages within the turbine section. A turbine stage may include a row of stationary airfoils, i.e., vanes, followed by a row of rotating airfoils, i.e., blades, where the turbine blades extract energy from the hot combustion gases for powering the compressor section and providing output power. Since the airfoils, i.e., vanes and blades, are directly exposed to the hot combustion gases, they are typically provided with an internal cooling passage that conducts a coolant, such as compressor bleed air, through the airfoil.
One type of turbine airfoil includes a radially extending outer wall made up of opposite pressure and suction sidewalls extending from leading to trailing edges of the airfoil. The cooling channel extends inside the airfoil between the pressure and suction sidewalls and conducts the cooling fluid in alternating radial directions through the airfoil.
In a turbine airfoil, achieving a high cooling efficiency based on the rate of heat transfer is a significant design consideration in order to minimize the volume of coolant air diverted from the compressor for cooling.